The present invention relates to an improvement in rocket motors and more particularly, but not by way of limitation, to a high performance, safe, non-polluting, thrust controllable liquid-solid propulsion system.
Rocket propulsion systems have found extensive application as assist take-off devices, boosters for missiles, projectiles and various spacecraft. The systems previously used may be classified into two basic types in accordance with the physical state of the fuel and oxidizer employed, i.e., liquid propellant or solid propellant. A third type, a gaseous system is undesirable because of the very large space requirements in order to provide an adequate fuel supply.
The liquid propellant systems have advantages in controllability, low cost, and high performance but all liquid oxidizing agents present severe storage and handling problems. For example, some liquid oxidizers are corrosive, unstable and toxic while others are toxic and volatile and others require insulation and refrigeration. In addition, conventional liquid systems require complex feed mechanisms and precise control to eliminate the hazard of explosion. The solid propellant systems are simpler and may be more reliable than the liquid propellant systems, but they are explosive, non-controllable and expensive.
Several systems employing fuels and oxidarts, respectively in different physical states, had been developed. One such system comprised the reaction of aluminum powder suspended in liquid hydrocarbon with liquid oxygen and involved the attendant disadvantages of the liquid type systems noted above. Another system proposed the use of liquid nitrous oxide for burning solid carbon and still another system proposed burning a solid plastic with a liquid oxidizer. The latter two systems also involved the problem of storing and handling a highly corrosive, unstable, or toxic liquid.
Solid oxidizers were developed to provide auxiliary, combustion-supporting agents for fuel rich solid propellants. These oxidizers were included in a rocket chamber containing a solid fuel, e.g., nitroglycerine or nitrocellulose, for supplying additional oxygen to promote combustion of the solid fuel and thereby increase the combustion efficiency. This was essentially a solid type system and embraced the disadvantageous features of non-controllability, expensiveness and short duration operation.
Certain prior art patents while displaying the shortcomings discussed are of interest. For example, U.S. Pat. No. 4,214,439 is of interest in that it shows a multi-component propulsion system utilizing metallic fuel particles and a carrier gas as a propellant medium. Liquid hydrogen is pumped through a fuel regenerative passage in the walls surrounding the rocket housing to act as a coolant for the combustion chamber walls and becomes vaporized during this cooling step. The vaporized hydrogen is then pumped into a fluidization chamber through a bed of metallic particles to maintain them in a fluidized condition. The fluldized particles and the hydrogen are then directed to the combustion chamber to be combined with a liquid oxidizer.
U.S. Pat. No. 3,136,119 discloses a fluid-solid propulsion unit wherein compressed air and liquid fuel are introduced into a combustion chamber lined with a hollow cylindrical grain of solid oxidizing material. The liquid fuel and compressed air are initially burned and applied to the solid grain until the intense heat produced causes the grain to decompose thereby evolving oxygen which reacts with more fuel to further intensify the heat released until the combustion of the solid grain becomes self sustaining. The supply of oxygen is then shut off.
U.S. Pat. No. 4,441,312 shows a combined cycle ramjet engine wherein the metallic wall of the combustion chamber of a combined rocket-ramjet engine is provided with solid ramjet fuel overlaid with rocket fuel. After the consumption of the rocket fuel in the boost portion of the flight, the solid ramjet fuel burns and ablates protecting the metallic combustion chamber wall from high temperatures during the cruise phase of the missile flight.
U.S. Pat. No. 3,844,118 shows a solid fuel ramjet powered missile having the air inlets positioned adjacent the aft end of the fuel grain with a channel provided adjacent the fuel grain to provided primary air into the forward end of the fuel grain with the secondary air being supplied directly to the secondary combustor. A valve is provided in the primary air flow channel to control the air flow into the fuel grain.
U.S. Pat. No. 3,073,110 shows a propellant utilization control system which controls the fuel and oxidizer flow to a sustainer engine of a rocket system so that both the propellant tank containing the liquid oxygen and the propellant tank containing the fuel are emptied at the same time. Utilizing a closed loop control system, the ratio of the oxidizer to the fuel supplied to the engine is maintained substantially equal to the ratio of the oxidizer to the fuel in the propellant tanks during the operating period of the engine system.
In light of the foregoing review of previous rocket propulsion system and illustrative patented systems and their attendant shortcomings the present invention has as a general object, the provision of a rocket propulsion system which combines the beneficial features of both the liquid and solid type systems but avoids the undesirable features of each.